Hiber Response to Di

REPLY submitted by Hiber Inc.

Hiber Response

2019-02-22

This document pretains to SAT-PDR-20180910-00069 for Petition for Declaratory Ruling on a Satellite Space Stations filing.

IBFS_SATPDR2018091000069_1628365

1 8 0 0 M S T R E E T, N W SU ITE 800N W ASHINGTON, DC 20036 TEL 202.783.4141 FA X 202.783.5851 W W W .W BKL AW .C O M February 22, 2019 Mr. Jose P. Albuquerque Chief, Satellite Division, International Bureau Federal Communications Commission 445 12th Street, S.W. Washington, DC 20010 Re: Hiber, Inc., Petition for Declaratory Ruling Access U.S. Market Using the Hiberband Low-Earth Orbit System; Call Sign S3038, IBFS File No. SAT-PDR-20180910-00069 Dear Mr. Albuquerque: Hiber Inc. (“Hiber”) hereby responds to the letter dated November 20, 2018 (“Division Letter”),1 from the Satellite Division (“Division”) of the International Bureau requesting additional information regarding Hiber’s Petition for Declaratory Ruling (“PDR”) seeking U.S. market access for a non-voice, non-geostationary (“NVNG”) mobile-satellite service (“MSS”) system in the 399.9-400.05 MHz and 400.15-401 MHz frequency bands (“Hiberband® System”).2 Specifically, Hiber provides answers to Questions 1-6 regarding its orbital debris analysis report (“ODAR”). Question 1: The Orbital Debris Assessment Report does not appear to be fully executed in the signature block. The version of the Debris Assessment Software utilized is not a current version. Please update. Response: Attached is a fully executed copy of Hiber’s updated ODAR. 1 See Letter from Jose P. Albuquerque, Chief, Satellite Div., Int’l Bur., FCC, to Lynne Montgomery, Wilkinson Barker Knauer, LLP, Counsel to Hiber, Inc., IBFS File No. SAT-PDR-20180910-00069, Call Sign S3038 (Nov. 20, 2018) (“Division Letter”). 2 See Hiber, Inc., Petition for Declaratory Ruling to Access U.S. Market Using the Hiberband Low-Earth Orbit System, IBFS File No. SAT-PDR-20180910-00069 (Call Sign S2979) (filed Sept. 10, 2018) (“PDR”).

Mr. Jose P. Albuquerque February 22, 2019 Page 2 As previously noted, Hiber has been unable to obtain a current version of the Debris Assessment Software 2.1.1 (“DAS”) from NASA. 3 Due to difficulty with processing pending and additional requests, NASA has been unable to provide a timeframe for providing Hiber with the current DAS version. Consequently, Division staff has indicated that alternative debris assessment software may be used. Hiber has chosen to use the National Centre for Space Studies’s (“CNES”) Semi-analytic Tool for End of Life Analysis (“STELA”) and the European Space Agency’s (“ESA”) Debris Risk Assessment and Mitigation Analysis (“DRAMA”) software applications to assess risk probabilities, as noted in the attached ODAR. Question 2: In its Orbital Debris Assessment Report, Hiber identifies two failure modes that may be inversely related. The first – “lithium plating on the anode” – is caused by operation below recommended temperatures, while the second – “gas generation” – is caused by use above recommended temperatures. Please provide recommended temperature range and any steps to avoid operations above or below this range. The recommended temperature ranges are set forth in Section 3.4 of the attached ODAR. In particular, the recommended temperature ranges for the battery cells are: • Charge : 0 to +45° C • Discharge : -20 to +60° C • Storage : -20 to +50° C Hiber will continuously monitor the battery cell temperatures on Hiber-1 and Hiber-2. To ensure that operations remain within the recommended temperature ranges, Hiber can turn the heaters on or off to adjust the battery cell temperature Question 3: In its Orbital Debris Assessment Report, no calculations or data are included to support Hiber’s conclusions regarding the probability of collision with space objects. Please provide additional information on these calculations. Hiber used ESA’s DRAMA software to assess the probability of collision with space objects in Section 3.5 of the attached ODAR. DRAMA allows Hiber to assess the compliance of its mission with international safety and debris requirements. DRAMA computes the annual collision probability. The probability of collision was calculated for each satellite’s orbital lifetime. 3 See Letter from Lynne Montgomery, Wilkinson Barker Knauer, LLP, Counsel to Hiber, Inc., to Jose P. Albuquerque, Chief, Satellite Div., Int’l Bur., FCC, IBFS File No. SAT-PDR-20180910-00069 (Call Sign S3038) (Dec. 12, 2018); Letter from Lynne Montgomery, Wilkinson Barker Knauer, LLP, Counsel to Hiber, Inc., to Jose P. Albuquerque, Chief, Satellite Div., Int’l Bur., FCC, IBFS File No. SAT-PDR- 20180910-00069 (Call Sign S3038) (Jan. 29, 2019) (“Jan. 29 Letter”); see also Division Grant Stamp.

Mr. Jose P. Albuquerque February 22, 2019 Page 3 To determine the evolution of the satellite’s orbital parameters, Hiber used CNES’ STELA software. STELA allows the user to propagate orbits over time and provides a complete report of the evolution of the orbital parameters throughout the satellite’s lifetime. Question 4: In its Orbital Debris Assessment Report, Hiber provides the probability of collision for a proposed satellite system of two satellites. Please provide orbital debris mitigation information for Hiber’s proposed twenty-four space station constellation. In Section 3.5 of the attached ODAR, Hiber calculated that the collision probability of an initial two-satellite constellation is 1.3485 x 10-4, which is lower than NASA’s 0.0001 threshold probability. 4 When considering the entire 24-satellite system (consisting of two 6U satellites and 22 3U satellites), the total collision probability is calculated as 2.2383 x 10-3, which exceeds the 0.001 threshold. 5 However, the 22 3U satellites are expected to be equipped with propulsion modules. The propulsion modules will enable Hiber to conduct collision avoidance maneuvers as needed. Question 5: The Schedule S lists the estimated lifetime of the satellites as three years from date of launch. There is no further documentation regarding the length of time these satellites will be in orbit through natural decay. Please provide additional information supporting Hiber’s conclusion regarding the lifetime of the satellites. This information should be provided showing altitude and time data band may be submitted in a graph format. Hiber performed simulations using CNES’ STELA software to assess how long it would take for the Hiber-1 and Hiber-2 satellites to effectuate an atmospheric reentry. The results show that the satellites will take 4.54 years and 16.74 years, respectively, to re-enter the atmosphere after the end useful life. This is compliant with the guidelines that specify that satellites de-orbiting through atmospheric reentry do so within 25 years of the satellite’s end-of-life. The results are demonstrated in Figures 1 and 2 below. 4 See NASA Technical Standard, Safety and Mission Assurance Acronyms, Abbreviations, and Definitions, NASA-STD 8709.22 at 32, Requirement 4.5-1 (with Change 2) (Oct. 31, 2012) (setting standard that the probability of a spacecraft colliding with a large object during the satellite’s orbital lifetime should be no greater than 0.001). 5 As noted in its application, Hiber is finalizing the design for the 22 3U satellites and will submit a separate ODAR when the design is complete.

Mr. Jose P. Albuquerque February 22, 2019 Page 4 Figure 1 : Hiber-1 altitude evolution Figure 2 : Hiber-2 altitude evolution Question 6: Section 3.7 of the Orbital Debris Assessment Report includes an incomplete table of spacecraft components. Please provide a complete list of spacecraft components.

Mr. Jose P. Albuquerque February 22, 2019 Page 5 Section 3.7 of the attached ODAR has been updated with a complete list of spacecraft components. Should the Commission require additional information about the foregoing or otherwise in connection with the PDR, please contact the undersigned. Sincerely, /s/ Lynne M. Montgomery Lynne M. Montgomery Counsel to Hiber, Inc. cc: Jose Albuquerque Karl Kensinger Stephen Duall Alyssa Roberts

hiber. Orbital Debris Assessment Report Hiber—1 & Hiber—2 Revision nr2 2210212018 This document contains no proprietary, ITAR or export controlled information Once this document has been printed it will be considered an uncontrolled document.

Revision nr2 | Page 2 Hiber Orbital Debris Assessment Report Name Prepared by Hiber Inc. Version Approved by Maarten Engelen;Program Executive/Project Manager Signature /// /’ Date February22,2019 s#o~" @Ghiber

Revision nr2 | Page 3 Revisions Revision Description Date 1 Initial release 16/08/2018 2 First revision 20/11/2018 3 Use of STELA and DRAMA instead of DAS 2.1.1 06/02/2018 Updated launch dates and orbit parameters of Hiber—1 and 2 Specified recommended temperature range for battery cells Completed the probability of collision chapter Computed time before reentry for Hiber—1 and 2 Completed table for spacecraft components Chiber.

Revision nr2 | Page 4 Table of Contents 7 RIHOUNUGHIOM :sss—:.sss—ssssrsssrinerrenrsinriaref¥ITREEYEAEETArEea n trv i rraververypanrte¥resvers sns ¥y¥re¥sYFrYrresrrraver eR es 248 5 3.1 PUFROSE ... . . ... . . .. . »»+ «o y eoupomnnt roemers rearereie ds 1 eaviaire in ns en y Bi‘e HEHERTEHAE ENE We s t uis se se ue ie uie e uie e wre B wrareran e avereren t 5 1.2 ROTOROMOOS:.. : . : : :s : :»=2 18 sme0 B e nE CERENERSATESE ts w CA 1 nie t it ield o ol ie e wl uie ape alea e is y Eies are B i aiuia‘aia n nleicia uin in arere e en e srare ie $ 2 GBNOFAILROVIQWWL——.s—srsssserrrrrerririrerrirrirrrristrtreryscress¥¥sFESSSEEEYEEEYEEEsrrrrTrrererFrerrFra Th cueirevvicesnr s E299 6 3 Orbital DebrIS ASSOSSMQONE .....s..ssssssssrssec.0ssrr es sns rersrirssrriivss ies rrrrrererr ie e trrerearerrrerserrersrere 7 3.1 Program Management and MiSSION OVE@rVI@W ........22222222222202220rrrrrvrrrvrrrrrrrrrrrrrrr e e ie e e ie e se e e k ies T 3.2 SDAGCGCFETE DOSGCFIQUHONY . ; . : :9222 8 208 0# m00000e 10 00+ mm on sn wis‘e n n ind n is e ce Wt ue w i a it n n i e a n se a e es en ie ale es e in i itc 7 3.8 Assessment of Spacecraft Debris Released during Normal Operations ....................c.clclll. 9 3.4 Assessment of Spacecraft Intentional Breakups and Potential for Explosions ...................... 9 3.5 Assessment of Spacecraft Potential for On—Orbit COIliSIONS.........2...2.02222200e e rrrerrrrr e rr e e 10 3.6 Assessment of Spacecraft Post—mission Disposal Plans and Procedures .......................... 13 3.7 Assessment of Spacecraft Re@ntry HAZArUS ........................0062¥¥¥¥¥¥r k¥ rrevrrrerrrrerrrerrrrerrereree . 155 3.8 Assessment of Spacecraft Hazardous MAteri@IS .........................00000000eeeererrrrrrrrrrrrrerrrr en 166 3.9 Assessment fOr Teth@r MISSIONS..........2.2222200222200002era¥rrerrrrrrrrrrrr en rrrrr en e e e e e rr e e e e en rr e e e en r e e e e e e e ed 16 List of Tables Table 1: List Of FeferenC@dIRNOCUMIEMES. . . . . . . . . :. . ... » «s cmmensnmie n in e meaiee selne s onl c diee se rivi en ue m me ue uie se taiala niaa‘s on aaini n 5 Table 2 : ODAR rSevVigw CNOCK SMNOSL. .. :. . . . . .. : . :: — : . .—sexserereeseeedrk s Fery HETEIETEEEAT men es neTt nembie ns ns e e sns wie n ue eiaiee ns se amreurenne 6 Table 3 : Battory TaIIUF®: MOUES. . ... . ... . .. . . . .. . . .0. 1. onecrnone re oo nied t ce e FEeeERET EiEA a‘e w hoi Pss Pe id on wie m nials e ue B es mee e aie ue aieioe ie 10 Table 4 ~Orbit parameters BVOIUTION . . . ... ... . . ; :. .: :s . :« cce eremmerinrersinende resseenenenee renmnee mm en en n ns n ns ces nisve‘s 11 Table 5 : Annual collision probaldllitiGs ...... ... ... ...... ... —on.orerserrerener ie xBrFerie es Feve e nere rri e rererrennecernen beveneiinpnes 12 Table 6 : Description of spacecraft COMPONEOMNES ................ :..1.rerrsrversassvrrnsraes inss se iss rnerernreneesaterrereneres 16 Table 7 : Hazardous materials fOUNd ON SDAC@CFAft...................22220000202rearrrrrrrrrrrrrrrrrrrrrrrrrrrrrrrrrrrrrrrrkee0 16 List of Figures FIgure 1 : Hibers SATOIIIIG . :: ::#3 :mme0rztensshies m 10 +s3 9i i s nini siteinlni sn ieia ie in vran e uie ie ie l uie ue Ei in Eie w o e a e in e e e uce e en oi 8 Figure 2 : Hiber—1 AltitUU@ @VOIUtION..........2.220220220eee e rrerrerrerrerrrrrrerreeees Error! Bookmark not defined.3 Figure 3 : Hiber—2 AIt{UOG BVOILUNIOM.. .. .. :11 .221 s 22sssssersss rrrsrskrrcerserercverreees Error! Bookmark not defined.4 hiber. _

Revision nr2 | Page 5 1 Introduction 1.1 Purpose The scope of this document is to assess the different causes and risks related to orbital debris. Only the risks related to Hiber‘s satellites, and not the launch vehicle, are being investigated in this report. This study was conducted in concordance with the requirements stated in NASA—STD—8719.14A. In order to determine risk probabilities, Hiber used the National Centre for Space Studies‘ (CNES) Semi—analytic Tool for End of Life Analysis (STELA) software and the European Space Agency‘s (ESA) Debris Risk Assessment and Mitigation Analysis (DRAMA) software. 1.2 References Ref # Document / software Version 1 NASA—STD—8719.14A 2 |STELA 3.2 3 |DRAMA 2.2.1 Table 1: List of referenced documents Chiber. —

Revision nr2 | Page 6 2 General Review The following table summarizes the different requirements recommended by NASA. As can be seen, Hiber‘s spacecraft meet all requirements. Launch Vehicle Spacecraft Regquirement 4 Not Stanhdard Compliant Not Comments # Compliant Compliant Incomplete ConN1;:;ant or N/A Compliant Incomplete No intentional 4.3—1 .a X X debrig released No intentional 4.3—1.b X X debris released No intentional 4.3—2 X X debris released 4,4—1 X X No 4.4—2 xX X passivation No 4.4—3 X X intentional breakup No 4. 4—4 X X intentional breakup 4.5—1 X X 4.5—2 X X 4.6—1 .a X X 4.6—1.b X X 4.6—1.¢c X X 4.6—2 X X 4.6—3 X X 4.6—4 X X 371 X X No tether 4.8—1 X * system Table 2 : ODAR review check sheet fl'hiberfi _»

Revision nr2 | Page 7 3 Orbital Debris Assessment 3.1 Program Management and Mission Overview * Identification of the Headquarters Mission Directorate sponsoring the mission and the Program Executive: The Headquarter Mission Directorate is Hiber and the Program Executive is Maarten Engelen. * Identification of the responsible program/project manager and senior scientific and management personnel: The responsible project managers are Tom Schreuder and Maarten Engelen. * Identification of any foreign government or space agency participation in the mission and a summary of NASA‘s responsibility under the governing agreement(s): None. * Clear schedule of mission design and development milestones from NASA mission selection through proposed launch date, including spacecraft PDR and CDR (or equivalent) dates: N/A * Brief description of the mission: The first two 6U satellites of the Hiber constellation, Hiber—1 and Hiber—2, have been deployed in a LEO orbit. Their mission lifetime will be 3 years. 22 additional satellites will be launched to form a 24—satellite constellation. These additional satellites will be 3U in size, and will be studied in a separate ODAR as their design will differ. * Identification of the anticipated launch vehicle and launch site: Hiber—1 was launched on a PSLV—C43 from the Satish Dhawan Space Centre (Sriharikota, India) and Hiber—2 on a Falcon 9 from the Vandenberg Air Force Base (California, USA). * Identification of the proposed launch date and mission duration: Hiber—1 was launched on November 29th, 2018 and Hiber—2 on December 3rd, 2018. Both satellites are designed for a mission duration of three years. * Description of the launch and deployment profile, including all parking, transfer, and operational orbits with apogee, perigee, and inclination: Hiber—1 was launched into an orbit with a perigee at 479.7 km, apogee at 507.6 km and a 97.5° inclination. Hiber—2 was launched into an orbit with a perigee at 580.3 km, apogee at 599.9 km and inclination at 97.8°. The satellites will then naturally decay because of atmospheric drag forces. * Reason for selection of operational orbit(s) (such as ground track, SSO, GEO sync, instrument resolution, co—locate with other spacecraft, ...): These orbits were chosen because they are sun—synchronous. The exact orbital planes are based on the placements by the launch vehicle providers. * Identification of any interaction or potential physical interference with other operational spacecraft (Note: This does not include potential for RF interaction unless it affects the risk of generating orbital debris.): None 3.2 Spacecraft Description * Physical description of the spacecraft, including spacecraft bus, payload instrumentation, and all appendages, such as solar arrays, antennas, and instrument or attitude control booms: Khiber. ——

Revision nr2 | Page 8 The Hiber—1 and Hiber—2 spacecraft are typical 6U CubeSats, with outside dimensions of 100 mm x 200 mm x 340.5 mm. There are two large solar panels (200 mm x 340.5 mm) and two smaller ones (100 mm x 340.5 mm), which will deploy once in orbit. Four antennas (550 mm long) serve for TTC, and two other antennas (one 330 mm long deployable antenna and one patch antenna) provide payload functionalities. * Detailed illustration of the entire spacecraft in the mission operation configuration with clear overall dimensional markings and marked internal component locations: 0.2 m CubeStar CubeSpace Battery Modulas ISIS Power Conditioning Controle Modules ISt Payload Ss Hyperion 3—Axis Contral System CubeSpace Power Distribution Model ISIS Medium Reaction Wheels TRXVU CubeSpace ISIS Payload Antenna CDHS and PDHU HCT ISIS TXS IS1S 0.330 m Figure 1 : Hiber‘s satellite * Total spacecraft mass at launch, including all propellants and fluids: 7.23 kg * Dry mass of spacecraft at launch, excluding solid rocket motor propellants: 7.23 kg * Description of all propulsion systems (cold gas, mono—propellant, bi—propellant, electric, nuclear): Hiber—1 and Hiber—2 do not include a propulsion system. * Identification, including mass and pressure, of all fluids (liquids and gases) planned to be on board and a description of the fluid loading plan or strategies, excluding fluids in sealed heat pipes. Description of all fluid systems, including size, type, and qualifications of fluid containers such as propellant and pressurization tanks, including pressurized batteries: N/A Khiber.

— Revision nr2 | Page 9 * Description of all active and/or passive attitude control systems with an indication of the normal attitude of the spacecraft with respect to the velocity vector: The satellite is controlled via 3 reaction wheels and 3 magnetorquers. The normal attitude of the spacecraft consists of the long axis nadir—aligned, as the payload antennas need to be nadir—pointing. This is the naturally stable attitude of the satellite. * Description of any range safety or other pyrotechnic devices: Hiber—1 and Hiber—2 do not contain any range safety or pyrotechnic devices. * Description of the electrical generation and storage system: Hiber—1 and Hiber—2 contain Li—ion battery cells that provide energy, and are recharged by GaAs solar cells. * Identification of any other sources of stored energy not noted above: There are no additional sources of stored energy. * Identification of any radioactive materials on board or make a positive statement that there are no radioactive materials onboard: There are no radioactive materials onboard. 3.3 Assessment ofSpacecraft Debris Released during Normal Operations No object will be released intentionally. 3.4 Assessment ofSpacecraft Intentional Breakups and Potentialfor Explosions * Identification of all potential causes of spacecraft breakup during deployment and mission Operations: There is no credible scenario which would lead to a spacecraft breakup. * Summary of failure modes and effects analyses of all credible failure modes which may lead to an accidental explosion: The only potential source of explosion on the satellite is from the Li—ion battery cells. Explosion can occur due to overheating or venting. Metal scraps that are left over from the manufacturing process, shocks, damages, over—discharging and overcharging, fast charging and using the batteries outside of the recommended temperature ranges could lead to an explosion. * Detailed plan for any designed spacecraft breakup, including explosions and intentional collisions: N/A *« List of components which are passivated at EOM. List includes method of passivation and amount which cannot be passivated: No items will be passivated. The batteries do not need to be passivated because, as the failure mode analysis shows, they do not present a high or credible risk of explosions during the mission. * Rationale for all items which are required to be passivated, but cannot be due to their Design: N/A — Assessment of spacecraft compliance with Requirements 4.4—1 through 4.4—4 Chniber.

Revision nr2 | Page 10 Requirement 4.4—1: Limiting the risk to other space systems from accidental explosions during deployment and mission operations while in orbit about Earth or the Moon: For each spacecraft and launch vehicle orbital stage employed for a mission, the program or project shall demonstrate, via failure mode and effects analyses or equivalent analyses, that the integrated probability of explosion for all credible failure modes of each spacecraft and launch vehicle is less than 0.001 (excluding small particle impacts) The following failure modes were assessed regarding the battery cells. They can all potentially lead to explosion. Effects of Failure mode Causes of failure Recommended action failure Short circuit Overheating Metal scraps, shock, Quality check, vibration test, shock test, physical damage, charge and discharge cycling tests, over—discharge, discharge and overcharge protection, short overcharge, external circuit protection on external circuits system failure Overcharging Overheating No overcharging Overcharging protection protection Charge cycling test Overpressure Venting Ultra—fast charging Nominal charging Lithium Physical Use below Maintain battery cells at recommended plating on damage / recommended temperatures anode venting / short temperatures circuit Gas Venting Use above Maintain battery cells at recommended generation recommended temperatures temperatures Table 3 : Battery failure modes The recommended actions listed above were taken during the manufacture and operation of the satellites, hence mitigating the risk of battery cell explosions. Through these actions, Hiber has effectively mitigated against the risk of failure. Note: The recommended temperature range for the battery cells are: e Charge: 0 to +45° C e Discharge: —20 to +60° C e Storage: —20 to +50° C The battery cell temperature is continuously monitored. Hiber can turn the heaters on or off to adjust the battery cell temperature to ensure that operations remain within the recommended temperature range. Requirement 4.4—2: Passivate to limit probability of accidental explosion after EOM: N/A Requirement 4.4—3: Limiting the long—term risk to other space systems from planned breakups: N/A Requirement 4.4—4; Limiting the short—term risk to other space systems from planned breakups: N/A 3.5 Assessment ofSpacecraft Potentialfor On—Orbit Collisions + Calculation of spacecraft probability of collision with space objects larger than 10 cm in diameter during the orbital lifetime of the spacecraft (compliance with requirement 4.5—1): hiber. __

Revision nr2 | Page 11 To determine the evolution of the satellite‘s orbit parameters, Hiber used CNES‘ STELA software. STELA allows the user to propagate orbits over time and provides a complete report of the evolution of the orbital parameters throughout the satellite‘s lifetime. Hiber used DRAMA to compute the annual collision probability. DRAMA allows Hiber to assess the compliance of its mission with international safety and debris requirements. The collision probability was calculated consistent with the orbital life of each satellite, each time with the updated orbit parameters. According to STELA, the evolution of the orbital parameters for Hiber‘s constellation is: Parameter Hiber—1 Hiber—2 Hiber 3—24 Time Launch date 11/29/18 12/03/18 10/01/19 Lifetime 4.54 years 16.74 years 4.58 years 3 Semi—major axis (km) 6871.65 6968.10 6978 5s Eccentricity 0.00125 0.00121 0.00122 S Inclination (°) 97.5 97.8 97.8 Arg. of perigee (°) 91.5 91 87.5 Semi—major axis (km) 6869.25 6967.62 6976.50 ‘c_‘B Eccentricity 0.00 1245 0.001205 0.001235 § Inclination (°) 97.54 97.84 97.79 Arg. of perigee (°) 91 91 86.5 oJ Semi—major axis (km) 6865.62 6966.89 6968.99 s Eccentricity 0.00 1235 0.001215 0.001255 $ Inclination (°) 97 °57. 97.87 97.77 Arg. of perigee (°) 90.5 91.25 87.5 & Semi—major axis (km) 6847.89 6963.04 6939.11 i Eccentricity 0.001185 0.00119 0.001325 $ Inclination(°) 97.60 97.91 97.76 Arg. of perigee (°) 90 91.25 88 is Semi—major axis (km) 6784.71 6952.99 6870.15 s Eccentricity 0.001145, 0.00118 0.001515 £ Inclination (°) 97.61 97.94 97.74 Arg. of perigee (°) 87.5 90.5 87.5 in Semi—major axis (km) 6940.79 s Eccentricity 0.00118 yo Inclination (°) 97 .97 Arg. of perigee (°) 90 5 Semi—major axis (km) 6930.81 iz Eccentricity 0.00120 y Inclination (°) 97.98 Arg. of perigee (°) 89.25 = Semi—major axis (km) 6925.26 is Eccentricity 0.00122 § Inclination (°) 97 .97 Arg. ofperigee (°) 88.75 3 Semi—major axis (km) 6922.20 is Eccentricity 0.00122 2 Inclination (°) 97.94 Arg. of perigee (°) 88.75 & Semi—major axis (km) 6920.51 s Eccentricity 0.001225 £ Inclination (°) 97.91 Arg. of perigee (°) 88.5 y & . Semi—major axis (km) 6919.28 £ Eccentricity 0.00123 ghiber@ _

— Revision nr2 | Page 12 Inclination (°) 97.87 Arg. of perigee (°) 88.25 & Semi—major axis (km) 6918.06 m Eccentricity 0.00123 $ Inclination (°) 97.85 l Arg. of perigee (°) 88.25 o Semi—major axis (km) 6915.94 > Eccentricity 0.00124 $ Inclination (°) 97.84 = Arg. of perigee (°) 88.25 53 Semi—major axis (km) 6908.09 &E Eccentricity 0.00127 $ Inclination (°) 97.85 & Arg. of perigee (°) 88 \f Semi—major axis (km) 6887.22 o. Eccentricity 0.001305 6 Inclination (°) 97 .87 = Arg. of perigee (°) 88 i Semi—major axis (km) 6851.48 s Eccentricity 0.00134 $ Inclination (°) 97.90 >— Arg. of perigee (°) 89 «3 Semi—major axis (km) 6787.42 C Eccentricity 0.00136 $ Inclination (°) 97.91 * Arg. of perigee (°) 90.75 Table 4 : Orbit parameters evolution The annual collision probabilities are given on DRAMA: Collision probability Hiber—1 Hiber—2 Hiber 3—24 Year 0—1 0.1513 x 10—4 0.2694 x 10—4 0.3815 x 10—4 Year 1—2 0.1849 x 10—4 0.3342 x 10—4 0.3238 x 10—4 Year 2—3 0.1727 x 10—4 0.3039 x 10—4 0.3279 x 10—4 Year 3—4 0.1482 x 10—4 0.2137 x 10—4 0.2180 x 10—4 Year 4—5 0.6072 x 10—5 0.2135 x 10—4 0.1603 x 10—4 Year 5—6 0.1948 x 10—4 Year 6—7 0.2120 x 10—4 Year 7—8 0.2071 x 10—4 Year 8—9 0.2094 x 10—4 Year 9—10 0.2340 x 10—4 Year 10—11 0.2031 x 10—4 Year 11—12 0.1966 x 10—4 Year 12—13 0.2132 x 10—4 Year 13—14 0.1979 x 10—4 Year 14—15 0.1932 x 10—4 Year 15—16 0.1488 x 10—4 Year 16—17 0.5802 x 10—5 Total 0.71782 x 10—4 3.60282 x 10—4 1.4115 x 10—4 Table 5 : Annual collision probabilities The total collision probability for both Hiber—1 and Hiber—2 is 4.32064 x 10—4, which is under the 0.001 threshold. When considering the 24 satellite system the total collision probability becomes 3.5374 x 10—3, thus exceeding the 0.001 threshold. However, the 22 3U satellites, for which Hiber will submit a separate niber. _

Revision nr2 | Page 13 ODAR, are expected to be equipped with propulsion modules. The propulsion modules will enable Hiber to conduct collision avoidance maneuvers as needed, thus mitigating the possibility of collision. * Calculation of spacecraft probability of collision with space objects, including orbital debris and meteoroids, of sufficient size to prevent post—mission disposal (compliance with requirement 4.5—2): Post—mission disposal is done naturally, via drag forces. Therefore, there are no vital systems needed to ensure it. Similarly, no systems will be passivated, so once again there will be no vital systems needed to ensure it. 3.6 Assessment ofSpacecraft Post—mission Disposal Plans and Procedures + Description of spacecraft disposal option selected: The spacecraft will decay because of atmospheric drag and de—orbit naturally via atmospheric re—entry. Simulations were run on CNES‘ STELA software to assess how long it would take for the satellites to effectuate an atmospheric reentry. Results show that it will take Hiber—1 4.54 years and Hiber—2 16.74 years to reenter the atmosphere. This is compliant with the guidelines that specify that satellites de—orbiting through atmospheric reentry do so within 25 years of the satellite‘s end—of—life. The results are demonstrated in Figures 2 and 3. Semi—major axis Nature: Mean parameters CNES — STELA version: 3.2 cars saso * meg seas 6800 e7rs 6750 s72s Semi—major axis (km) 6700 5675 a a o 2 6sas 6600 6575 esso 6525 6500 o a.s 1 1.5 2 2.5 3 25 4 4.5 Time (years) — semi—major axts | Figure 2: Hiber—1 altitude evolution flhiben 2 (

Revision nr2 | Page 14 Semi—major axis Nature: Mean parameters CNES — STELA version: 3.2 cors eoso t Hiirreaged c925 sns eon0 ~— cors x 6850 \ 6925 @ 6800 \\\\\ Semi—major axis (km 6775 @ @S $ oN g si 2 u $ ho f 3 oon a a & h ow N P 6525 ssoo 6575 6550 eS25 6500 0 2.5 s 7.5 10 12.5 is 17.5 Time (years) [Semi—major is | Figure 3: Hiber—2 altitude evolution * Identification of all systems or components required to accomplish any postmission disposal operation, including passivation and maneuvering: N/A *« Plan for any spacecraft maneuvers required to accomplish postmission disposal: N/A + Calculation of area—to—mass ratio after postmission disposal, if the controlled reentry option is not selected: The mass of the satellite will be 7.23 kg at the end of life. At the end of life, the satellite will no longer be controlled and will automatically align its longest axis with nadir. This means the cross—sectional area will be either 0.03405 mor 0.0681 m, depending on which side is oriented with the velocity axis. As a result, the area to mass ratio will vary between 4.71 x 10° m*/kg and 9.42 x 10° m*/kg. * If appropriate, preliminary plan for spacecraft controlled reentry: N/A + Assessment of spacecraft compliance with Requirements 4.6—1 through 4.6—4 Requirement 4.6—1 : Disposal for space structures passing through LEO: Compliant. The satellites will reenter the atmosphere within 25 years. Requirement 4.6—2 : Disposal for space structure near GEO: N/A Requirement 4.6—3 : Disposal for space structures between LEO and GEO: N/A Requirement 4.6—4 : Reliability of post—mission disposal operations in earth orbit: Because the disposal operation is natural and will happen automatically, it is entirely reliable.

Revision nr2 | Page 15 3.7 Assessment ofSpacecraft Reentry Hazards * Detailed description of spacecraft components by size, mass, material, shape, and original location on the space vehicle, if the atmospheric reentry option is selected Diam / : Group Name Quantity Material Shape to Mass J width 1# Eength i Helght EPS PCCM + 1 FR4 Box 0.429 0.100 0.100 0.050 PDM (fiberglass) BM 3 Li—ion Box 0.220 0.100 0.100 0.033 _depSPA s 2 GaAs Box 0.279 0.100 0.3405 0.00015 depSPA l 2 GaAs Plate 0.557 0.200 0.3405 0.00015 CDHS iOBC 1 FR4 Box 0.096 0.100 0.100 0.014 PDHS PDHU 1 FRA4 Box 0.099 0.100 0.100 0.014 TXS—2 1 FR4 Box 0.068 0.090 0.096 0.033 S—Patch 1 Ceramic Cylinder 0.420 0.084 x 0.0 148 TTG TRXVU 1 FRA4 Box 0.077 0.090 0.096 0.015 ANTs + 1 Aluminum Box 0.204 0.098 0.098 0.007 cover plate AOCS RW 3 Brass Box 0.153 0.046 0.046 0.0315 ADCS 1 FR4 Box 0.278 0.090 0.096 0.075 Board STR 3 FR4 Box 0.067 0.035 0.050 0.100 CSS 10 FR4 Box 0.015 0.001 0.003 0.001 FSS 2 FR4 Box 0.08 0.09 0.096 0.01 MTM 1 FR4 Box 0.015 0.03 0.03 0.005 MTM (red) 1 FR4 Box 0.071 0.03 0.03 0.1 Payload HCT 1 NiTinol Cylinder 0.290 0.100 0.100 0.330 BEE I FR4 Box 0.500 0.100 0.100 0.100 RO4350 GPS ant 1 Ceramic Box 0.040 0.035 0.035 0.0055 MECH Structure 1 Aluminum Box 1155 0.100 0.200 0.3405 6061 MISC 1GIS 1 FR4 Box 0.105 0.100 0.100 0.01 Harness 1 Copper Box 0.500 0.100 0.100 0.006 Table 6 : Description of spacecraft components Location: All components are inside the spacecraft, except for the antennas, solar arrays, sun sensors, and magnetometers. + Summary of objects expected to survive an uncontrolled reentry, using NASA Debris Assessment Software (DAS), NASA Object Reentry Survival Analysis Tool (ORSAT), or comparable software: Using ESA‘s DRAMA software, Hiber concludes that the ceramic S—Band and GPS antennas are expected to survive re—entry. The probability that both fragments reach ground is 93.041%. The probability that at least one fragment reaches the ground is 99.875%. The estimated casualty cross section area is 0.8024 m2 and the estimated mass at ground impact is 0.415 kg. * Calculation of probability of human casualty for the expected year of uncontrolled reentry and the spacecraft orbital inclination (compliance with requirement 4.7—1): Using ESA‘s DRAMA software, the probability of human casualty is 9.3378 x 10°. The casualty probability is compliant with requirement 4.7—1, which requires a probability lower than 1/10 000. ghibefl /

Revision nr2 | Page 16 3.8 Assessment ofSpacecraft Hazardous Materials + Summary of the hazardous materials contained on the spacecraft: a Estimated i a Estimated Estimated Chemical Description state. Estimated Estimated state, state, and ial of how it is quantity, stat:.t, stattgt, quantity, quantity, zgm;ngfi;?e a hazard to activity, g;‘:s“s:l}é g:j:sns:lre’ pressure at pressure material humans pressure on orbit at EOM end of to survive at launch passivation reentry Li—ion Toxic gases Solid Solid Solid Solid None battery cell released 0.66 kg 0.66 kg 0.66 kg 0.66 kg Panasonic when NCR18650A exploding Table 7: Hazardous materials found on spacecraft 3.9 Assessmentfor Tether Missions There are no tether systems in the mission. Khiber. .


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Document Modified: 2019-02-22 16:38:00
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